In fixed wing and rotary wing such as helicopters, a pilot causes the Aircraft to undergo flight maneuvers via a hydraulic actuator. Typically, the pilot input forces are multiplied or boosted that provides a relatively large mechanical advantage to the pilot when altering flight control surfaces or rotor blade pitch. For example, as the pilot provides a small amount of force (e.g., about 0.5 pounds force) on the mechanical linkages, the hydraulic boost device provides a relatively large force (e.g., 2,000-5,000 pounds force) on the flight control surface or rotor blade.
Conventional hydraulic boost devices are designed to avoid flight-critical failure modes. Typically in helicopter or aircraft systems, multiple hydraulic actuators are structurally or mechanically linked together to provide redundancy. Such redundancy allows for continued safe flight of the helicopter in the event that a portion of a hydraulic system driving a hydraulic actuator fails during operation.
FIG. 1 illustrates one type of hydraulic actuator system 10 utilized to provide dual redundancy for flight controls. As illustrated, the hydraulic actuator system 10 includes two separate hydraulic cylinders 12, 14 positioned in either a side-by-side arrangement, as shown, or may also be in an inline tandem arrangement. A dedicated control valve 16 and hydraulic system 20 controls the first hydraulic cylinder 12 while a dedicated control valve 18 and hydraulic system 22 controls the second hydraulic cylinder 14. Both the control valve and the cylinder are structurally connected and perform identical functions for dual redundancy. In use, as a pilot actuates a lever 24, the lever 24 positions a valve member 26 of the first and second control valves 16, 18 to control the flow of pressurized fluid from the hydraulic systems 20 and 22 to the hydraulic cylinders 12, 14, thereby controlling the positioning of a flight control surface via the rod end 28. In the event one of one of the hydraulic systems 20, or 22 were to develop a leak or pump failure and system pressure were lost, the remaining functioning control valve and hydraulic system provides pressurized fluid to the corresponding hydraulic cylinder to allow for continued operation. The use of a dual and independent hydraulic actuator system 10 provides for a relatively high degree of safety by allowing the pilot to control the aircraft in the event of failure of a portion of the hydraulic actuator system 10.
FIG. 2 illustrates another type of hydraulic actuator system 50 utilized to provide dual redundancy. As illustrated, the hydraulic actuator system 50 includes a single hydraulic cylinder 52 with a single control valve 54. The hydraulic actuator system 50 also includes first and second hydraulic systems 56, 58 in fluid communication with the hydraulic cylinder 52 via a switching valve 60. The switching valve 60 provides failover redundancy in the hydraulic actuator system 50. For example, the switching valve 60 is disposed in a first compressed position relative to the control valve 54, as illustrated, to allow pressurized fluid to flow between the first hydraulic system 56 and the hydraulic cylinder 52. As a pilot actuates a lever 62, the lever 62 adjusts a position of a valve member 64 of the control valve 54 to control the flow of pressurized fluid from the first hydraulic system 56 to the hydraulic cylinder 52. In turn, the hydraulic cylinder 52 controls the positioning of rod end 68. In the event that the hydraulic system 56 were to develop a failure, the switching valve 60 moves to a second, extended position via spring 70 to allow pressurized fluid to flow between the second hydraulic system 58 and the hydraulic cylinder 52. The use of the single control valve 54 in combination with the switching valve 60 provides the system 50 with a level of redundancy at a relatively low cost and low weight.